Rotationally Adaptive Flight Test Surface

Final Report.... /(0

1......... i'

by

Ron Barrett

Assistant Professor of Aerospace Engineering

Aubum University, AL 36849

(334) 844-6825 (office) -6832 (lab) -6803 (FAX) rbarrett@eng.aubum.edu

for

Mr. David Voracek

National Aeronautics and Space Administration

Dryden Flight Research Center

Warehouse #6, Bldg. 4876

Edwards, California 93524

(805) 258-2463

under contract number: NCC4-103

15 April 1999

Abstract

Research on a new design of flutter exciter vane using adaptive materials was

conducted. This novel design is based on all-moving aerodynamic surface technology

and consists of a structurally stiff main spar, a series of piezoelectric actuator elements

and an aerodynamic shell which is pivoted around the main spar. The work was built

upon the current missile-type all-moving surface designs and change them so they are

better suited for flutter excitation through the transonic flight regime. The first portion of

research will be centered on aerodynamic and structural modeling of the system. USAF

DatCom and vortex lattice codes was used to capture the fundamental aerodynamics of

the vane. Finite element codes and laminated plate theory and virtual work analyses will

be used to structurally model the aerodynamic vane and wing tip. Following the basic

modeling, a flutter test vane was designed. Each component within the structure was

designed to meet the design loads. After the design loads are met, then the deflections

will be maximized and the internal structure will be laid out. In addition to the structure, a

basic electrical control network will be designed which will be capable of driving a

scaled exciter vane. The third and final stage of main investigation involved the

fabrication of a 1/4 scale vane. This scaled vane was used to verify kinematics and

structural mechanics theories on all-moving actuation. Following assembly, a series of

bench tests was conducted to determine frequency response, electrical characteristics,

mechanical and kinematic properties. Test results indicate peak-to-peak deflections of

1.1 deg with a corner frequency of just over 130 Hz.

1. Background of the Problem or Opportunity' "

During the past decade, great strides have been made in the area of flight control

through adaptive materials. Many of these systems have been centered on shape-

memory alloy, piezoceramic and magnetostrictive actuation of flight control surfaces.

Some of the earliest groundbreaking work was performed on active vibration control. 1

These early studies were rapidly followed by innovations in modeling which

demonstrated that relatively simple laminated plate theories could be used to accurately

model the performance of many types of adaptive structures.2, 3 These studies paved

the way for attempts at flight control. Through aeroelastic effects, it was thought that

small active deflections could be magnified to produce usable deflections. 4-6 At about

the same time as the early active aeroelastic estimations, it was discovered that twist

could be directly imparted to uncoupled structures by piezoceramic elements. This

discovery of directionally attached piezoelectric (DAP) elements opened the door to a

host of aerodynamic control devices. 7-13

Modifications to the conventional DAP structures were made to include

aeroservoelastic deformations. Several studies showed that small deflections could be

effectively magnified just like earlier conventional adaptive structures approaches. 14

One of the first application-oriented studies used a new concept dubbed active free-spar

torque-plate missile fins. These novel fins employed the torque plates and DAP

elements of earlier studies, but, for the first time, they were applied to practical flight

vehicles. 15-17 These torque-plate fins provided moderate deflections with moderate

forces, but for some flight regimes, smaller pitching moments are common. In particular,

the low subsonic flight regime yields nearly negligible pitching moments about the

aerodynamic center. Accordingly, another family of adaptive aerostructure actuators

was developed. These Flexspar actuators provided the highest recorded active pitch

deflections for flight control purposes. 18-21

Although fine for the low subsonic flight regime, these actuators very quickly become

saturated by pitching moments at transonic Mach numbers. The center of pressure shift

simply overpowers the relatively weak adaptive structures within the shell. To counter

this adverse property, yet another class of (unpublished) actuators were developed

specifically for the transonic flight regime. These rotationally active linear actuators

(RALA) are capable of utilizing up to 95% of the active strain energy of the

piezoceramics (rather than the 10 to 40% which is captured by other methods).

Because the strain energy capture rate is so much higher, the moment-deflection work

zones are nearly an order of magnitude greater than preceding actuators. As with all

new approaches, however, a penalty is paid in total deflection angle. Accordingly, these

new actuators are high moment, low displacement devices. 22

As the development of active aerostructures has progressed, new innovations in the

flutter flight test devices have also made great strides. Numerous flutter exciters which

are based on inertial, aerodynamic and pyrotechnic forces have been fabricated and

used. 23-27 One of the most modem devices circumvents many of the difficulties which is

found with each of the earlier systems. This rotating cylinder-based structural excitation

system has been shown to provide adequate levels of excitation at the desired

frequency ranges and moderate power consumption levels. 28 However, there exists an

opportunity to use the adaptive aerodynamic flight control device technology to improve

upon the performance levels which were obtained by the rotating cylinder-based

structural excitation system.

Because adaptive elements are much more efficient than conventional

electromechanical actuators, the power consumption of the system will be significantly

reduced. Also, the solid state nature of such actuators generally widens the frequency

response of the system. An added benefit which is important to specific structural

excitation schemes is the ability to use non-sinusoidal and superimposed waveforms.

The non-sinusoidal waveform input will place pseudo-step force inputs into the aircraft.

Such a square-wave excitation scheme will tend to induce higher structural modes. The

superimposed waveform capability will allow for testing structural interactions and cross-

couplings. Such tests may be conducted with a steady excitation of one structural mode

(second bending for example) while simultaneously performing a frequency sweep in

the vicinity of another mode (first torsion for example). Because the system is capable

of step and superimposed responses, a white noise test signal may be fed to the fin.

Such a white noise signal may be used in conjunction with atmospheric turbulence to

drive high structural modes. Another benefit of such a system is that it will occupy no

volume within the wing and power leads may be run along the external wing surfaces by

electrically conducting tape. Such an easy installation means that multiple exciters may

be quickly applied to hard points on test aircraft. By using multiple adaptive vanes, they

may be driven by the same or different signals; but, more importantly, they may be

phased so as to leverage excitation levels. Clearly, the development of such an

advanced system will greatly enhance flutter testing through structural excitation.

Accordingly, this study is centered on investigating the fundamental properties of this

rotationally-active flutter test surface (RAFTS).

2. Technical Ob ectives

This investigation is centered on demonstrating the basic feasibility of the RAFTS

exciter. As a first step, this study will establish the working principles of the RALA

actuators within a scaled flutter test surface. As a funding option, a full-scale test article

may also be constructed.

Table 2.1 Goals of RAFTS Project

i. Develop aerodynamic and structural models for RAFTS actuator vane

ii. Construct and bench-test a prototype RAFTS actuator vane for model verification

The first technical goal is set forth so that a new class of actuator vanes using the

RALA actuators may be developed. To accomplish this, a series of vortex-lattice codes

will be constructed for aerodynamic predictions. USAF DatCom methods will also be

used for modeling under separated flow conditions. The structural models will be built

upon laminated plate and virtual work theories as well as finite-element methods.

The second major goal is to bring a feasible RAFTS design to life. This new class of

exciter vane will be first modeled on a 1/4 scale for verification of the structural models.

A series of bench tests will confirm frequency response, moment generation capability

and deflection levels. A funding option will be used for the construction of a full-scale

exciter vane.

The first three months have yielded progress in two main areas on the

Rotationally Adaptive Flutter Test Surface (RAFTS). The first area of advance is in

aerodynamic modeling of the fin and selection of the preferred configuration. Thesecond area involves structural modeling of rotational actuators and design of the activemembers.

5

3. Aerodynamic and Structural Modeling

3.1. Aerodynamic Modeling of Fin Candidates and Selection of Preferred Fin

ConfigurationsAs laid out in the kick-off meeting of 3/96, one of the major goals of the project is

to maintain a level of control authority which is similar to existing flutter test surfaces.

Accordingly, a fin with similar control capability was designed. By using the Nsf

requirements, Ref. 29 and Ref. 30, it was possible to arrive at a range of fin designs

which would satisfy the normal force requirements. Because this work is initially being

performed on a 1/2 scale vane, the normal force estimations were cut in proportion tothe ratio of the areas (a factor of 4). Figure 1 shows the range of fin configurations

which provide control increments which are similar to that those for the rotating-cylinderbased structural excitation system (Ref. 3) at Mach 0.8 through 1.3. Because a second

major design condition is the maintenance of low pitching moments, the maximum

pitching moments were estimated for each design from Mach 0.8 to Mach 1.3. Thesemoments were optimized by determining the hinge location which minimizes moments

from 0 to 20 ° angle of attack. Figure 2 shows the results of the hinge moment

determination for a given increment of lift slope. Cleady, the higher the aspect ratio, thehigher the Na and the lower the Ma. Accordingly, a faidy high aspect ratio airfoil was

chosen with the limiting factor coming from aeroelastic and root bending moment

considerations. Clearly, the main spar of this section should be formed from high

strength composites or machined from high strength steel. Figure 3 shows the current

configuration and outer mold lines for the RAFTS vane as obtained from aerodynamic

optimization and force matching.

3.2. Structural Modeling of Rotational ActuatorsA series of basic structural deformation estimations were derived from laminated

plate theory and structural kinematics. By using the RALA configuration set forth

Section 3 of the proposal, an expression for the cross-sectional area of the individualactuator was derived:

A= (1)EL^2

Where A is the cross-sectional area, o is the maximum rotation angle desired, Eis the modulus of the active materials, L is the length of the material and A is the free-

strain of the actuator strip. The rotator rod diameter is determined by Equation 2:

d= 2LA (2)

6

With these estimations, assuming piezoceramic lead zirconate titanate, PZT-5H

at up to 220 pstrain over a 4" strip with +1 ° of rotation, the rotator rod diameter is

estimated to be 0.10" and the thickness of the active material strips is estimated at

0.033". Using 7.5 mil sheets to keep actuation voltages low, 5 layers of PZT will be

required per actuator for the 50% scaled RAFTS surface.

3.3. Aerodynamic Force MatchingFrom the aerodynamic data presented in Ref. 29, it can be seen that the dynamic

force in the transonic regime ranges from 52 to 65 Ib peak-to-peak at 30,000 ff altitude.By using the aerodynamic data of Ref. 30 and 31, a map of minimum normal forceslopes as a function of geometry may be obtained. If the geometries of Fig. 3.1 are usedas a survey of candidate shapes, then their geometric parameters and aerodynamiccharacteristics through the transonic flight regime may be obtained:

l]

31

3233

36

Figure 3.1 Candidate Vane Planforms

7

Table3.1Summaryof LowAspectRatioAerodynamicSurfaceDataI MachConfig. CNcm_n EquivalentFull-Span _.

No. (_^-1)0.0225

Aspect Ratio

I

0.8 to 1.3t_CLMmax

II I 0.044

14 0.0275 I 0 0.03

15 0.03 I 0.5 0.07

31 0.02 0.5 0.5 0.03

32 0.02 0.5 I 0.018

33 0.049 2 I 0.058

34 0.045 2 0.5 0.055

35 0.04 2 0 0.051

36 0.015 0.5 0 0.027

By using the above data, and assuming that the vane will be designed togenerate peak-to-peak normal force variations of at least 52 Ib at 30,000 ft, Mach 0.8,(282 psf dynamic pressure), the size and deflection range may be determined. From thedata of Ref. 29, Fig. 12, it can be seen that the greatest transonic demands on thesystem in terms of normal force/dynamic pressure occur at Mach 0.8. Accordingly, theRAFTS vane actuator will be sized at this most stringent condition.

3.4. Size and Deflection Determination

To approximate the force levels generated by the DEI exciter vane, a lft 2 area

was selected as the initial starting reference point as it was approximately the same

area as the DEI vane. By using the above information and assuming the 282 psf

dynamic pressure, and taking the minimum value for Cn(x (to be conservative), the

deflection levels of the various vanes can be estimated. Table 3.2 summarizes the

changes in angle of attack changes required for such a fin to generate normal forcevariations on the order of +26 Ibf at Mach 0.8, 30,000 ft.

Table 3.2 Summary of Minimum Aerodynamic Control and Geometric Parameters,

Math 0.8 - 1.3

Config. No. Na(deg ^-1) Aa (l/2p-p)(deg) Cr(in) MaximumAM(in-lb0

11 6.345 4.10 17 211

14 7.755 3.35 34 287

15 8.46 3.07 22.6 446

31 5.64 4.61 32 270

32 5.64 4.61 24 122

33 13.818 1.88 12 196

34 12.69 2.05 15.6 243

35 11.28 2.30 24 345

36 4.23 6.15 48 365

From Table 3.2, it is clear that configuration #33 requires the least pitch change

to generate the required +26 Ibf of normal force. Also, it has the second-lowest

maximum change in total pitching moment from 0 to 20 °, Mach 0.8 to Mach 1.3.

Because of these characteristics, this simple configuration was selected as the

configuration of choice.

8

The next logical step in the design of the vane is to choose the pivot point. To do

this, a detailed look at the shift in center of pressure through the transonic regime is

needed. From Ref. 30, a plot of c.p. migration from 0 ° - 12 ° at Mach 0.8 - 1.3 can be

examined. Table 3.3 shows the changes experienced in center of pressure location as afunction of angle of attack and Mach number.

Table 3.3 Chordwlse Center of Pressure Locatlon (o/_) as a Function of

Alpha (deg)

Mach an( Angle of Attack for Rn Conflguratlon #33 (Squa.85 1.lO

'e)1.20.80 .92 .98 1.30

.00

.00 .450 .450 .450 .450 .450 .450 .450

4.00 .183 .200 .193 .258 .286 .300 .331

8.00 .206 .197 .219 .302 .350 .359 .364

12.00 .298 265 .272 .330 .373 .372 .381

16.00 .367 .376 .378 .341 .377 .374 .398

By using the data of Table 3.3, and considering non-zero angles of attack, the

location of the optimum hinge line may be determined. By examining the difference

between the maximum and minimum pitching moment coefficients, a plot showing this

change, ACLM, can illustrate the region where the lowest moment requirements occur,

as shown in Figure 3.2.

0.18 -

0 I I I +

0.2 0.22 0.24 0.26 0.28 0.3 0.32 0.34 0.36 0.38 0.4 0.42

Hinge Line Reference Point, Xhl (%c)

0.44 0.46

Figure 3.2 Maximum Change in Hinge Moment Coefficient in the Transonic Regime

as a Function of Hinge Line Reference Point, Xhl (%c) for Fin Configuration #33

(Square)

9

From Fig. 3.2, it can be seen that there is a strong minimum in hinge moment

coefficient with a 40%c hinge line in the high angle of attack condition. Because,

however, the vane will generally be operating below 8 ° angle of attack, it is obvious that

a hinge location from 23%c to 34%c has approximately the same net change in pitchingmoment coefficient. Accordingly, the hinge line will be placed at 33%c to keep the low

angle of attach hinge moments low while approaching the minimum for the high angle ofattack regime.

To match the hinge moments which the vane will encounter, the RALA actuators

must be structurally balanced; accordingly, the actuator root width, Lr is one of the

prime factors which determines the performance of the actuators. Considering a pair of

0.030" thick PZT actuators used top and bottom, with strain actuation capability rangingfrom -300 to 200pstrain, a rotation-deflection plot may be generated. Because the

steady pitching moments are expected to range up to 120 Ibf, it is clear that low-force,

highly leveraged actuators will not be suitable. Figure 3 shows a map of actuator design

)arameters versus moment and deflection capability.

200 -

180

Eo5

60-

PZT-5H, -50°F to 120°F, 200 to -300 pstrain, 9msi

o. " 0.15" 0.10" 0.075" Lr = 0.05"

0 0.5 1 1.5 2 2.5 3

Rotational Deflection, e (deg)

3.5 4

Figure 3.3 Rotational Deflection and Active Moments Generated by a

0.030" Thick RALA Actuator Configuration

]0

From Fig. 3.3, considering the large pitching moments which the vane will

experience, it is obvious that the root length should be 0.20" or greater. Unfortunately,

this comparatively large dimension will effectively degrade deflections and accordingly,

unsteady force levels. However, there are two mitigating factors to this seemingly

discouraging result: 1) the -300 pstrain actuation level is approximately 1/2 of the

maximum which may ultimately be obtained from PZT-5H, and 2), because RAFTS may

be phased with respect to each other, their comparatively low force levels can be

effectively magnified when multiples of the vanes are used in unison. One should note

that the reason that the PZT-5H is not generally operated at the 600pstrain level is

because of difficulties with arcing and cracking. However, it is highly likely that with

better manufacturing techniques and stoichiometries of PZT, these limits may be safelycrossed with little fear of structural failure.

Even with those mitigating factors, the large steady force levels will still pose

problems for the RAFTS using current generation piezoceramics. In short, if the vanes

were designed to generate large deflections, the strong, steady upwash or downwash

field at the wing tip would simply drive the vanes against the rotation stops. There is,

however, a configuration which would practically eliminate this problem. If the vane

were allowed to rotate into the direction of the steady upwash field, then oscillate about

the zero angle of attack range (instead of some comparatively large angle of attack),

then the steady forces would be dropped by two or three orders of magnitude. To do

this, a simple weathercock vane should be added to the system behind the front vane.

Because the pivot is located at the 33%c, the most adverse nose-up moments will be

generated in the high subsonic flight regime where the dynamic pressure is high, and

yet the aerodynamic center is still located at the quarter-chord.

If a small 2" chord by 6" span vane is placed behind the forward vane, then the

aerodynamic center of the entire system will be shifted aft. Using a main pivot location

of 33%c with a 1 ft 2 forward vane followed by a 0.0833 ft 2 rear vane, the length of the

rear vane level arm may be solved for by taking a very conservative system

aerodynamic center of 50%c as exhibited by equation 1.

X,c = X=_f$:C_ + X--=_/S,CL,,, (3)s:cL +s,cL ,

where:

Sf = 1 ft 2 Sr = 0.08333 _ CL,_ = 0.049deg 1 CL,,, = 0.073deg 1

._,,,¢ = 0.33 '_oc = 0.50

The aerodynamic center location of the rear fin comes to:

X,,, = 2.5 X,,c, = 30in

11

With such an arrangement, the actuator system will possess positive stability

throughout the flight envelope by at least 17%. Such a large static margin guarantees

that the system will always rotate into the mean local flow field. From examination of theresults of Ref, 29, it can be seen that the local flow field varies widely up to 15 ° at the

wing tip (no doubt due to tip-vortex formation). Accordingly, the system stops must be

placed far enough apart so as to permit the vane assembly to freely rotate up 15°. If it

proves necessary in later tests, the stops in the assembly may be widened further.

Appendix A shows the overall layout of the RAFTS system including the forward andrear weathercock vanes and the root mount.

3.5. Internal Structure Layout

The intemal structure of the RAFTS forward adaptive vane is extremely simple. It

consists of a pair of PZT sheets bonded to the main spar in the front and to each otherin the back. The shell is connected to both points on the adaptive elements. These are

clearly shown in Appendix A.

12

4. Fabrication and Testing

4.1. Mold Components and Vane Shell FabricationThe components of the aluminum shell mold were completed and assembled.

Following cutting, the mold was processed to provide a suitable release surfaceconsisting of 2 rail thick Teflon sheets. More than two dozen different ply sequenceswere attempted till a suitable combination was achieved. Included in the lay-upsequence is a fiberglass protective outer layer, a lightning strike mesh, epoxy tape andthe underlying graphite main structural skin. Enclosed figures in Appendix A detail moreinternal components of the shells.

4.2. Main Spar and Root Mount Fixtures

The second main group of components includes the base block, main spar, and

mounting bar. These components were fabricated from aluminum and steel and formthe structural core of the RAFTS vane. Enclosed figures show more detail on these

components.

4.3. Tail Boom and Weathercock Vane

A graphite-epoxy composite tail boom and weathercock vane were alsofabricated and bonded to the aft section of the RAFTS mounting bar. These

components were co-cured in 350°F cures with Cyanimide 123 epoxy tape for superior

strength and durability. Attached figures lay out more details of these structures.

4.4. Piezoceramic Actuators and Accelerometers

The final group of components to be fabricated is the piezoceramic actuator. This

actuator is being constructed from PZT-5H sheets and will be matched to the

anticipated amplifier, once available. In addition to integrating the PZT actuators intothe shell structure, a pair of accelerometers are being integrated into the leading-edge

of the RAFTS vane to provide pitch and bending information during flight. Attached

figures in Appendix A show the overall geometry of the actuators.

4.5. Surface Work on the RAFTS Graphite ShellThe graphite shell on the RAFTS vane was taken from an unfinished state

through Deflashing to a dimensionally conformal shape. Figure 4.1 shows the RAFTSshell just after removal from the molds.

13

Figure 4.1 RAFTS Shell Immediately Following Removal from Composite Tools,Prior to Deflashlng

4.6. Integration of Main Spar and Piezoceramic Actuator SheetsThe steel spar was fitted to the piezoceramic actuator sheets in anticipation of

the amplifier. Final assembly of the actuator element will have to wait till the amplifierbecomes available so as to match dynamic impedance and mount only the number ofelements which the amplifier is capable of driving. Towards this end, the first layers ofpiezoceramic actuator sheets were bonded to the substrate and main spar. Figure 4.2shows the lay-up of the piezoceramic actuator sheets on the steel substrate.

Figure 4.2 Integration of Piezoceramic Sheets, Steel Substrate and Main Spar

14

4.7. Surface Finishing of the Weathercock Vane and Tail BoomThe surface finish of the weathercock vane and tail boom were worked on as well

as a fit check. Figure 4.3 shows the structures after bonding, deflashing and surfacefinishing.

Figure 4.3 Weathercock Vane and Tail Boom Assembly

4.8 Testing

The RAFTS vane was structurally checked to 200 Ibf of normal force with no sign

of yielding. The load was applied to the main spar at the 14% semispan from the root

fixture. Static testing showed 1.1 deg vane deflections with a corner frequency of 136

Hz and a natural frequency of 119 Hz as shown in Appendix B. Testing was executed

with the use of a sine signal generator driving an ACX piezoelectric amplifier.

Deflections were measured by Hall effect sensor and a National Instruments Data

Acquisition System and checked by using laser reflection techniques.

15

5. Ground Test Instructions

5.1 Physical DescriptionThe Rotationally Adaptive Flight Test Surface (RAFTS) vane is composed of

several distinct groups of components. Descriptions of the components along with

dimensions are contained in the attached pages. The base block acts as a mount for

the entire structure and should be firmly attached to the airframe or test fixture. The six

leads exit in the vicinity of the main spar bushing and should not be mechanicallystressed.

Within the base block is a main spar bushing and steel main spar. These items

firmly anchor the vane assembly to the base block. The main spar is attached to the

piezoceramic actuator sheets via an internal insert and epoxy bond. The piezoceramicsheets are sandwiched in two stacks of four 7.5 mil thick PZT-5H sheets. The aft portion

of the sheets is bonded to the aft of the RAFTS vane shell. The 30% chord of the shell

is connected to the main spar through a Teflon-slip coating on the main spar itself.

Although restrained from translating in any direction, the shell is able to freely rotateabout the main spar as commanded by the piezoceramic actuators. At the base of the

entire assembly is the vane mounting bar. This structure allows the vane too be pitchedinto the local freestream via the weather cock vane and tail boom assembly.

Instrumentation on the vane consists of two accelerometers mounted at the root

and tip of the front portion of the forward vane bay. The main spar also bears strain

gages at the root for arrangement in a typical cantilevered beam bridge.

5.2 Electrical ConnectionsA total of six wires exit the RAFTS vane. The forward blue wires bear

accelerometer signals. The remaining four carry signals to and from the strain gages

and piezoelectric actuators. These are shown schematically in Appendix C. They arelisted below:

1. Blue outer accelerometer cable

This cable carries signals from the accelerometer mounted on the outer tip of theRAFTS vane.

2. Blue wire inner accelerometer

This cable carries signals from the accelerometer mounted on the inboard cornerof the RAFTS vane.

3. Blue banded cable - two conductors

This cable carries two conductors with a grounded shielding. The blue conductor

is attached to the top strain gage. The white conductor is attached to the bottom strain

gage.

16

4. Blue banded cable - single conductor

This cable carries one conductor with a grounded shielding. The single conductoris connected to the center contact of both strain gages.

5. Red banded cable - two conductors

This cable carries two conductors with a grounded shielding. The two conductors

have been soldered together and should be operated and tested in this condition. Theinner contacts of the actuator stacks are connected to the two conductors. The reason

why there are two conductors and not just one, is that for repoling operations, the leads

will be desoldered and repoling fields may be applied independently to each of the

stacks, i.e. it is a feature for maintenance and high-level testing.

6. Red banded cable - single conductor

This cable carries one conductor with a grounded shielding. The single conductor

is connected to the outer contacts of both actuator stacks. Although this conductor is not

grounded, it is closest in physical proximity to the main spar which is indeed grounded.If at all possible, this single conductor should also be grounded.

5.3 Operational Limits and Procedures

It is recommended that the PZT actuators be powered by sine waves at all times.

If sine waves are unavailable, then duty-cycle balanced triangle waves will suffice.

Square waves are to be avoided at all times. The actuation connections should bemade between leads no. 5 and 6. The attached connection schematic shows the

voltage and frequency range which is recommended. Static voltage limits should be set

at +115 VDC. These limits hold through 5 Hz. As the frequency is increased, the voltage

should drop to ±50V at 100 Hz. No actuation is recommended beyond 200 Hz.

5.4 Testing

Deflections may be checked by using laser reflection techniques off the shell.

Calibrated performance plots are also attached for both static and dynamicperformance.

]7

Referencesii|

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13. Chen, P. and Chopra, I., "Feasibility Study to Build a Smart Rotor: Induced Strain Actuationof Airfoil Twist," From the Proceedings of the 1993 North American Conference on SmartMaterials and Structures, Albuquerque, NM.

14. Barrett, R., "Modeling Techniques and Design Principles of a Low Aspect Ratio ActiveAeroservoelastic Wing," Proceedings of the North American Conference on Smart Materialsand Structures, Albuquerque, New Mexico, pp. 107 - 118, 1993.

15. Barrett, R., "Active Composite Torque-Plate Fins for Subsonic Missiles," paper presented atthe Dynamic Response of Composite Structures Conference, New Orleans, Louisiana,August 30 - September 1, 1993.

16. Barrett, R., "Advanced Low-Cost Smart Missile Fin Technology Evaluation," Final Reportto Wright Laboratory, USAF Armament Directorate, contract number F08630-93-C-0039Eglin AFB, December, 1993.

18

17. Barrett, R., "All-Moving Active Aerodynamic Surface Research," presented at the 31stAnnual Technical Meeting of the Society of Engineering Science, College Station, TX,October 10- 12, 1994.

18. Barrett, R., "A Solid State Apparatus for Controlling Pitch Deflections of AerodynamicFlight Control Surfaces," Auburn Univ. invention disclosure, October, 1994 (patent pending).

19. Barrett, R., Gross, R. S., and Brozoski, F., "Missile Flight Control using Active FlexsparActuators," proceedings of the 1995 Smart Structures and Materials Conference, February26 - March 3, San Diego, California.

20. Barrett, R., Gross, R. S., and Brozoski, F. T., "Design and Testing of Subsonic All-MovingSmart Flight Control Surfaces," proceedings of the 36th AIAA Structures, StructuralD_ru_nics and Control Conference, New Orleans, LA, April, 1995, pp. 2289 - 2296, AIAApaper no. AIAA-95-1081.

21. Barrett, R., "Aeroservoelastic Characteristics of All-Moving Adaptive Flight ControlSurfaces," paper presented at the 10th VPI & SU symposium on Dynamics and Control, 8 -10 May, 1995.

22. Barrett, R., "Rotational Active Linear Actuators," Barrett Aerospace TechnologiesInvention Disclosure and Laboratory Report, September, 1995.

23. Kehoe, Michael W., "Aircraft Flight Flutter Testing at the NASA Ames-Dryden FlightResearch Facility," NASA TM-100417, May 1988.

24. Norton, William J., "Random Air Turbulence as a Flutter Excitation Source,"Proceedings of the 20th Annual Symposium, Society of Flight Test Engineers, Reno,Nevada, 1989, pp. 6.4-1 -6.4-11.

25. Goldman, A., Rider, C.P., and Piperias, P., "Flutter Investigations on a TransaviaPL12/T400 Aircraft," Aeronautical Research Laboratories, Melbourne, Australia, July1989.

26. Van Nunen, J.W.G., and Piazzoli, G., "Volume 9: Aeroelastic Flight Test Techniquesand Instrumentation," AGARD Flight Test Instrumentation Series, AG°1606, Feb.1979.

27. Larue, P., Millet, M., and Piazzoli, G., "Pyrotechnic Bonkers for Structural Tests inFlight," La Recherche Aerospatiale, May-June 1974, pp. 137 - 146.

28. Vernon, Lura, "In-flight Investigation of a Rotating Cylinder-Based StructuralExcitation System for Flutter Testing," NASA Technical Memorandum 4512, NASADryden Flight Research Facility, Edwards, Califomia, 1993.

29. Mottinger, T. A., and Washington, W. D., "An Experimental Investigation of Low

Aspect Ratio Fins Tested at Transonic Speeds on a Reflecting Plane," Aeroballistics

Directorate, U.S. Army Missile Command, Redstone Arsenal, Alabama, Report No.

RD-TM-71-15, July 1971.

30. Roskam, J., "Preliminary Calculation of Aerodynamic Characteristics," Airplane

Design, published by Roskam Aviation and Engineering Corporation, Ottawa, KS1990.

31. Vemon, L., "In-Flight Investigation of a Rotating Cylinder-Based Structural Excitation

System for Flutter Testing," NASA TM 4512, 1993.

19

12. References

1. Bailey, T. and Hubbard, J. E., " Distributed Piezoelectric-Polymer Active Vibration Control ofa Cantilever Beam," Journal of Guidance, Control and Dynamics, Vol. 8, No. 5, pp. 605-611, 1985.

2. Crawley, E. F. and De Luis, Javier, "Use of Piezoelectric Actuators as Elements of IntelligentStructures," A/AA Journal, Vol. 25, No. 10, Oct. 1987, pp. 987-997.

3. Crawley, E.F., Lazarus, K.B., and Warkentin, D. J., "Embedded Actuation and Processing inIntelligent Materials," 2nd Int. Workshop on Comp. Mat'Is and Struct. Troy, NY.

4. Lazarus, K. B., Crawley, E. F., Bohlmann, J. D., "Static Aeroelastic Control Using StrainActuated Adaptive Structures," proceedings of the First Joint U.S./Japan Conference onAdaptive Structures, Maui, Hawaii, October, 1990.

5. Song, O., Librescu, L. and Rogers, C. A., "Static Aeroelasticity Behavior of Adaptive AircraftWing Structures Modeled as Composite Thin-Walled Beams," presented at the InternationalForum on Aeroelasticity and Structural Dynamics, Aachen, Germany, June, 1991.

6. Ehlers, S. M., "Aeroelastic Behavior of an Adaptive Lifting Surface," Ph.D. Thesis, PurdueUniversity, August, 1991.

7. Barrett, R. M., "Intelligent Rotor Blade and Structures Development Using DirectionallyAttached Piezoelectric Crystals," M.S. Thesis, University of Maryland, College Park, MD,May, 1990.

8. Barrett, R. M., "Intelligent Rotor Blade Actuation through Directionally AttachedPiezoelectric Crystals," Presented at the AHS National Forum, Washington, D.C., May 1990.

9. Barrett, R. M., "Method and Apparatus for Structural Actuation and Sensing in a Desired' Direction," U. S. Patent Application 485,599/07, 1990.

10. Barrett, R., "Active Plate and Missile Wing Development Using EDAP Elements," Journalof Smart Materials and Structures, Institute of Physics Publishing, Ltd., Techno House,Bristol, UK, Vol.1, No. 3, pp. 214-226, ISSN 096.

1 1. Barrett, R., "Active Plate and Missile Wing Development Using DAP Elements," AIAAJournal, March, 1994.

12. Barrett, R., "Aeroservoelastic DAP Missile Fin Development," Journal of Smart Materialsand Structures, Institute of Physics Publishing, Ltd., Techno House, Bristol, UK, Vol. 2, No.2, pp. 55-65, ISSN 0964-1726, June 1993.

13. Chen, P. and Chopra, I., "Feasibility Study to Build a Smart Rotor: Induced Strain Actuationof Airfoil Twist," From the Proceedings of the 1993 North American Conference on SmartMaterials and Structures, Albuquerque, NM.

14. Barrett, R., "Modeling Techniques and Design Principles of a Low Aspect Ratio ActiveAeroservoelastic Wing," Proceedings of the North American Conference on Smart Materialsand Structures, Albuquerque, New Mexico, pp. 107 - 118, 1993.

15. Barrett, R., "Active Composite Torque-Plate Fins for Subsonic Missiles," paper presented atthe Dynamic Response of Composite Structures Conference, New Orleans, Louisiana,August 30 - September 1, 1993.

16. Barrett, R., "Advanced Low-Cost Smart Missile Fin Technology Evaluation," Final Reportto Wright Laboratory, USAF Armament Directorate, contract number F08630-93-C-0039Eglin AFB, December, 1993.

17. Barrett, R., "All-Moving Active Aerodynamic Surface Research," presented at the 31stAnnual Technical Meeting of the Society of Engineering Science, College Station, TX,October 10- 12, 1994.

18. Barrett, R., "A Solid State Apparatus for Controlling Pitch Deflections of AerodynamicFlight Control Surfaces," Auburn Univ. invention disclosure, October, 1994 (patent pending).

19. Barrett, R., Gross, R. S., and Brozoski, F., "Missile Flight Control using Active FlexsparActuators," proceedings of the 1995 Smart Structures and Materials Conference, February26 - March 3, San Diego, California.

20. Barrett, tL, Gross, R. S., and Brozoski, F. T., "Design and Testing of Subsonic All-MovingSmart Flight Control Surfaces," proceedings of the 36th AIAA Structures, StructuralDynamics and Control Conference, New Orleans, LA, April, 1995, pp. 2289 - 2296, AIAA

paper no. AIAA-95-1081.

21. Barrett, R., "Aeroservoelastic Characteristics of All-Moving Adaptive Hight Control

Surfaces," paper presented at the 10th VPI & SU symposium on Dynamics and Control, 8 -l0 May, 1995.

22. Barrett, R., "Rotational Active Linear Actuators," Barrett Aerospace TechnologiesInvention Disclosure and Laboratory Report, September, 1995.

23. Kehoe, Michael W., "Aircraft Flight Flutter Testing at the NASA Ames-Dryden FlightResearch Facility," NASA TM-100417, May 1988.

24. Norton, William J., "Random Air Turbulence as a Flutter Excitation Source,"Proceedings of the 20th Annual Symposium, Society of Flight Test Engineers, Reno,Nevada, 1989, pp. 6.4-1 -6.4-11.

25. Goldman, A., Rider, C.P., and Piperias, P., "Flutter Investigations on a TransaviaPL12/T400 Aircraft," Aeronautical Research Laboratories, Melbourne, Australia, July1989.

26. Van Nunen, J.W.G., and Piazzoli, G., "Volume 9: Aeroelastic Flight Test Techniquesand Instrumentation," AGARD Flight Test Instrumentation Series, AG-1606, Feb.1979.

27. Larue, P., Millet, M., and Piazzoli, G., "Pyrotechnic Bonkers for Structural Tests inFlight," La Recherche Aerospatiale, May-June 1974, pp. 137 - 146.

28. Vernon, Lure, "In-flight Investigation of a Rotating Cylinder-Based StructuralExcitation System for Flutter Testing," NASA Technical Memorandum 4512, NASADryden Flight Research Facility, Edwards, Califomia, 1993.

29. Mottinger, T. A., and Washington, W. D., "An Experimental Investigation of Low

Aspect Ratio Fins Tested at Transonic Speeds on a Reflecting Plane," Aeroballistics

Directorate, U.S. Army Missile Command, Redstone Arsenal, Alabama, Report No. RD-

TM-71-15, July 1971.

30. Roskam, J., "Preliminary Calculation of Aerodynamic Characteristics," Airplane

Design, published by Roskam Aviation and Engineering Corporation, Ottawa, KS 1990.

31. Vernon, L., "In-Flight Investigation of a Rotating Cylinder-Based Structural Excitation

System for Flutter Testing," NASA TM 4512, 1993.

Appendix A ,

RAFTS Vane Figures

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Appendix B

RAFTS Test Data

1.6

Static IL_,FTS Actuator Tip Pitch Deflection_Jth Shell Removed

._ i.4r,

•_ _1.2

;-__o.6- _0.4.ag

_ o.2

0 20 40 60 80 100 120

1/2 Peak-to-Peak Applied Voltage, V (VDC)

Static RAFTS Vane Pitch Deflection

0

0 20 40 60 80 100 120

1/2 Peak-to-Peak Applied Voltage, V (VDC)

Dynamic RAFTS Vane Pitch Deflection

2.5

m

o 2

_v_._ 1.5

" 0.5

0 20 40 60 80 100 120 140 160

Excitation Frequency. f (Hz)

= ,

Appendix C

RAFTS Connection Schematic

Connection Schematicfor RAFTS Vane

Cable 1 Blue

Outer Accelerometer

Cable 2 Blue

Inner Accelerometer

Cable 3 Blue banded two conductor cable

outer strain gage contacts

((

120 top strain oao_

.-,,_'_ 120 bottom strain gage,,--"v'Vx,--'-

(_

Cable 4 Blue banded single conductor cable

inner center strain gage contact

Cable 5 Red banded two conductor cable

outer strain gage contacts

connected for normal operatIon. I hey

are separate only for maintenance and

_ advanced testing purposes.

top PZT actuator (eulward poling)

I

Cable 6 Red banded single conductor cable

inner center strain gage contact

...5ff

Q) C'l

•*-' O-8 ""> '

v

Recommended operational range

operationa_

/ "_ _..- _----__

00 5 10 20 50 100 200 500

frequency, .f (Hz)